This application was made with government support under Contract No. N00019-02-C-3003 awarded by the United States Navy. The Government may therefore have certain rights in this invention.
This application relates to turbine vane cooling.
Gas turbine engines typically include a compression section which compresses air. The compressed air is mixed with fuel and combusted in a combustion section. Products of that combustion pass downstream over turbine rotors, which are driven to rotate. The turbine rotors carry blades, and typically have several stages. Stationary vanes are positioned intermediate the stages. The stationary vanes are subject to extremely high temperatures from the products of combustion. Thus, cooling schemes are utilized to provide cooling air to the vanes.
A vane typically includes an airfoil and intermediate platforms at each end of the airfoil. It is known to provide platform cooling holes. In general, the vanes have been cast as a thin wall generally hollow item at their platform, and cooling holes have been drilled through the thin wall.
While the cooling holes provide some modest level of film cooling to the vane platforms, as temperatures of combustion increase, it would be desirable to provide both a more uniform and increased level of cooling effectiveness along the platform surface.
It becomes desirable to incorporate a cooling scheme that provides both active backside convective cooling along with more effective gas path film cooling.
It is known to provide a teardrop shaped cooling feature at the trailing edge of the airfoil. A teardrop shape cooling feature has a shape defined by flow dividers with a shape that is generally similar to a teardrop, and results in certain flow characteristics. However these features have not been used to facilitate film cooling along other high heat load regions of the airfoil and platform surfaces.